Composite mandrel for autoclave curing applications

ABSTRACT

A composite mandrel includes a generally elongated mandrel body comprising a resilient mandrel core and an elastomeric mandrel outer layer disposed outside the mandrel core. A method for fabricating a contoured stiffened composite panel is also disclosed.

TECHNICAL FIELD OF THE INVENTION

The disclosure relates to mandrels for forming cavities in compositematerials. More particularly, the disclosure relates to a compositemandrel which is suitable for autoclave curing applications in theformation of cavities in composite materials.

BACKGROUND OF THE INVENTION

When composite materials are molded into shapes with cavities, such ashat stringers, for example, there may be a need for some type of toolingthat can apply pressure from the cavity outward during the curing stepand can be extracted from the cavity after curing. The existing toolingused for this purpose may include without limitation inflatable rubbermandrels; solid mandrels such as metal, rubber or composite mandrels; ordissolvable mandrels. However, the inflatable rubber mandrels may beprone to leaking, which may lead to widespread porosity in the resultingcomposite laminate. The solid rubber mandrel may result in a cavity witha distorted cross-sectional shape or exert an uneven pressure on thecomposite laminate and may be too heavy for fabrication of large parts.The solid metal or composite mandrels may not have sufficientflexibility to be removed from parts having any degree of curvature orcomplexity. The dissolvable mandrels may be expensive to make anddifficult to remove from large parts. Existing mandrel designs may notaccommodate the dimensional changes of the composite part which occursduring application of heat to the surrounding tooling and part materialsat the curing step. This can cause undesirable part material movementresulting in such distortions as waviness, wrinkling and/or bridging inthe composite material.

Therefore, a mandrel is needed which is suitable for curing applicationsin the formation of cavities in composite materials and overcomes someor all of the limitations of conventional composite mandrels.

SUMMARY OF THE INVENTION

The disclosure is generally directed to a composite mandrel. Anillustrative embodiment of the composite mandrel includes a generallyelongated mandrel body comprising a resilient mandrel core and anelastomeric mandrel outer layer disposed outside the mandrel core. Themandrel may combine the desired characteristics of foam and rubber toproduce a manufacturing aid for airplane stringers or other similar opencavity parts made from fiber/resin composite materials. Themanufacturing aid which is embodied in the composite mandrel may be lesscostly, more durable and less prone to failures than current inflatablebladder technologies.

BRIEF DESCRIPTION OF THE ILLUSTRATIONS

FIG. 1 is a top view of an illustrative embodiment of the compositemandrel.

FIG. 2 is a cross-sectional view, taken along section lines 2-2 in FIG.1, of the composite mandrel.

FIG. 3 is a cross-sectional view of an alternative illustrativeembodiment of the composite mandrel.

FIG. 4 is an exploded top view of a composite assembly, moreparticularly illustrating insertion of multiple composite mandrels intorespective stiffening elements in the composite assembly preparatory tocuring of the composite assembly.

FIG. 5 is a cross-sectional view, taken along section lines 5-5 in FIG.4, of the composite assembly.

FIG. 6 is a top view of the composite assembly, with the compositemandrels inserted in the respective stiffening elements of the assembly.

FIG. 7 is a top view of the composite assembly, contained in vacuumbagging preparatory to curing of the assembly.

FIG. 8 is an exploded top view of the composite assembly, moreparticularly illustrating removal of the composite mandrels from therespective stiffening elements in the composite assembly after curing ofthe composite assembly.

FIG. 9 is a flow diagram which illustrates an illustrative method forfabricating a contoured stiffened composite panel.

FIG. 10 is a flow diagram of an aircraft production and servicemethodology.

FIG. 11 is a block diagram of an aircraft.

DETAILED DESCRIPTION

Referring initially to FIGS. 1 and 2, an illustrative embodiment of thecomposite mandrel is generally indicated by reference numeral 1. Thecomposite mandrel 1 may be used to fill a cavity (not shown) in anairplane stringer or other open-cavity part (not shown) made fromfiber/resin composite materials to prevent collapse of the cavity duringcuring of the composite materials. The composite mandrel 1 may be lesscostly, more durable and more effective and reliable than currentinflatable bladder mandrel technologies.

The composite mandrel 1 includes a generally elongated mandrel body 7having a mandrel core 2 which is a resilient material and a mandrelouter layer 10 which is disposed outside the mandrel core 2, as shown inFIG. 2, and is an elastomeric material. In some embodiments, the mandrelcore 2 is foam or other such material which incorporates open spaceand/or air pockets to prevent bulk modulus behavior during thermalexpansion and the mandrel outer layer 10 may be an elastomeric materialsuch as elastic rubber, for example and without limitation. The mandrelcore 2 and the mandrel outer layer 10 may be generally coextensive withthe mandrel body 7.

The mandrel core 2 and the mandrel outer layer 10 may have anycross-sectional shape depending on the particular use requirements ofthe composite mandrel 1. In some applications, for example, each ofmultiple composite mandrels 1 may be suitably configured to fillrespective stiffening elements (such as stringers) 27 during the curingand/or cocuring of a composite panel assembly 24, as shown in FIGS. 4-8and will be hereinafter described. As shown in FIG. 2, in someembodiments of the composite mandrel 1, the mandrel body 7 may have agenerally triangular cross-sectional shape. Accordingly, the mandrelcore 2 has a generally flat or planar core base 3 with lateral coreedges 6. Core sides 4 angle from the respective core edges 6. A coreapex 5, which may be rounded, extends between the core sides 4. Theshape of the mandrel outer layer 10 may generally correspond to that ofthe mandrel core 2, defining a mandrel base 11 which extends adjacent tothe core base 3; a pair of mandrel sides 12 which extend adjacent to therespective core sides 4; a mandrel apex 13 which may be rounded and isdisposed adjacent to the core apex 5; and mandrel edges 14 whichcorrespond positionally to the respective core edges 6 of the mandrelcore 2.

As shown in FIG. 3, in some embodiments of the composite mandrel 1 a,the mandrel body 7 a may have a generally trapezoidal shape.Accordingly, the mandrel core 2 a has a generally flat or planar corebase 3; a pair of core sides 4 which angle from the core base 3; and agenerally flat or planar mandrel core top 8 which extends between thecore sides 4. The mandrel outer layer 10 a defines a mandrel base 11which extends adjacent to the core base 3; a pair of mandrel sides 12which extend adjacent to the respective core sides 4; a generally flator planar mandrel top surface 16 which is disposed adjacent to themandrel core top 8; and mandrel edges 14 which correspond to therespective core edges 6 of the mandrel core 2 a.

Referring next to FIGS. 4-8, in typical application, multiple compositemandrels 1 are inserted in respective stiffening elements 27 provided ina stiffening layer 26 of a composite panel assembly 24 during curing ofthe composite panel assembly 24. The composite panel assembly 24 willultimately form an airplane stringer (not shown); however, it will beappreciated by those skilled in the art that the composite mandrels 1can be adapted to fill cavities in any other type of open-cavity orclosed-cavity composite material part made from fiber/resin compositematerials during curing of the composite material part. The compositemandrels 1 can be adapted to fill cavities having a constantcross-sectional shape or a cross-sectional shape which varies along thelength of the composite material, such as cavities which taper or curvealong the length of the cavity, for example and without limitation.

As illustrated in FIG. 5, in an embodiment of fabrication of thecomposite panel assembly 24, a base composite layer 25 may initially beplaced on a tooling surface 20 of OML tooling or IML tooling, forexample and without limitation. The tooling surface 20 may have agenerally concave contour, as shown. Alternatively, the tooling surface20 may have a generally planar or convex contour, depending on theparticular application. The stiffening layer 26 may be placed on thebase composite layer 25. The stiffening elements 27 may be shaped in thestiffening layer 26 and extend along the longitudinal axis of thetooling surface 20 in generally parallel relationship with respect toeach other, as shown in FIG. 4, and in generally perpendicularrelationship with respect to the concave contour of the tooling surface20. Alternatively, the stiffening elements 27 may be separate ordiscrete units. As further shown in FIG. 5, each stiffening element 27has a stiffening element cavity 28. In some embodiments, the stiffeningelements 27 may be oriented in orientations other than along thelongitudinal axis of the tooling surface 20 and may converge or diverge,for example and without limitation.

As shown in FIGS. 4 and 6, multiple composite mandrels 1 may be insertedinto the stiffening element cavitys 28 of the respective stiffeningelements 27. The elastomeric mandrel outer layer 10 of each compositemandrel 1 allows for a proper fit of the composite mandrel 1 into thestiffening element cavity 28 of each stiffening element 27 and conformsto pad-ups and ramps. As shown in FIG. 7, the composite panel assembly24 may then be enclosed in vacuum bagging 30 and cured by autoclaving.During the curing process, the composite mandrels 1 maintain the shapeand prevent collapse of the respective stiffening elements 27 as thecomposite material of the base composite layer 25 and the stiffeninglayer 26 hardens.

After curing, the composite panel assembly 24 is removed from the vacuumbagging 30. The composite mandrels 1 may be removed from the stiffeningelement cavitys 28 of the respective stiffening elements 27, as shown inFIG. 8. During removal, the elastomeric mandrel outer layer 10 of eachcomposite mandrel 1 may easily be deformed; this reduces the effortrequired for removal. The cured composite panel assembly 24 may then beprocessed to complete fabrication of the airplane assembly (not shown)or other composite part, according to the knowledge of those skilled inthe art.

It will be appreciated by those skilled in the art that the resilientmandrel core 2 of the composite mandrel 1 enhances the structural andcompressive characteristics of the composite mandrel 1 relative to thedesigns of conventional mandrels. This structural and compressivesupport may be necessary to maintain the shape of the stringer or othercomposite part during automated composite fiber placement as well asautoclave curing. Since the outer mandrel layer 10 may be a constantthickness, it may expand uniformly during curing, thus avoiding theproblems associated with uneven expansion of a solid rubber material.The cross-sectional area and type of foam used for the mandrel core 2may be engineered to impart compression compliance under autoclavepressure, thus offsetting the combined thermal expansion behavior of thefoam and rubber.

Referring next to FIG. 9 of the drawings, a flow diagram 900 whichillustrates an illustrative method for fabricating a contoured stiffenedcomposite panel is shown. In block 902, a tooling surface, such as thetooling surface 20 which was heretofore described with respect to FIG.5, for example and without limitation, is provided. The tooling surfacemay have a concave, planar, convex or alternative contour. In block 904,a base composite layer is laminated on the tooling surface. In block906, open-section stiffening elements are positioned on the basecomposite layer. In block 908, composite mandrels are provided. Eachcomposite mandrel includes a resilient mandrel core and an elastomericmandrel outer layer disposed outside the resilient mandrel core. Inblock 910, composite mandrels are inserted in the respective stiffeningelements. In block 912, the composite panel and stiffening elements aresealed in vacuum bagging. In block 914, the composite panel and thestiffening elements are cured. An autoclave may be used during curing.In block 916, the composite mandrels are removed from the stiffeningelements.

Referring next to FIGS. 10 and 11, embodiments of the disclosure may beused in the context of an aircraft manufacturing and service method 78as shown in FIG. 10 and an aircraft 94 as shown in FIG. 11. Duringpre-production, exemplary method 78 may include specification and design80 of the aircraft 94 and material procurement 82. During production,component and subassembly manufacturing 84 and system integration 86 ofthe aircraft 94 takes place. Thereafter, the aircraft 94 may go throughcertification and delivery 88 in order to be placed in service 90. Whilein service by a customer, the aircraft 94 may be scheduled for routinemaintenance and service 92 (which may also include modification,reconfiguration, refurbishment, and so on).

Each of the processes of method 78 may be performed or carried out by asystem integrator, a third party, and/or an operator (e.g., a customer).For the purposes of this description, a system integrator may includewithout limitation any number of aircraft manufacturers and major-systemsubcontractors; a third party may include without limitation any numberof vendors, subcontractors, and suppliers; and an operator may be anairline, leasing company, military entity, service organization, and soon.

As shown in FIG. 11, the aircraft 94 produced by exemplary method 78 mayinclude an airframe 98 with a plurality of systems 96 and an interior100. Examples of high-level systems 96 include one or more of apropulsion system 102, an electrical system 104, a hydraulic system 106,and an environmental system 108. Any number of other systems may beincluded. Although an aerospace example is shown, the principles of theinvention may be applied to other industries, such as the automotiveindustry.

The apparatus embodied herein may be employed during any one or more ofthe stages of the production and service method 78. For example,components or subassemblies corresponding to production process 84 maybe fabricated or manufactured in a manner similar to components orsubassemblies produced while the aircraft 94 is in service. Also, one ormore apparatus embodiments may be utilized during the production stages84 and 86, for example, by substantially expediting assembly of orreducing the cost of an aircraft 94. Similarly, one or more apparatusembodiments may be utilized while the aircraft 94 is in service, forexample and without limitation, to maintenance and service 92.

Although the embodiments of this disclosure have been described withrespect to certain exemplary embodiments, it is to be understood thatthe specific embodiments are for purposes of illustration and notlimitation, as other variations will occur to those of skill in the art.

What is claimed is:
 1. A method for fabricating a contoured stiffenedcomposite panel for an aircraft structure, comprising: placing a basecomposite layer on a tooling surface; placing at least one stiffeningelement having a stiffening element cavity on said base composite layer;inserting a one-piece resilient mandrel body in said stiffening elementcavity, the one-piece resilient mandrel body comprising a foam core andan elastomeric outer layer substantially co-extensive with the foamcore, and wherein the one-piece resilient mandrel body substantiallyfills the cavity of the stiffening element; wherein a cross-sectionalarea and type of foam used for the foam core is engineered to impartcompression compliance under autoclave pressure to offset a combinedthermal expansion behavior of the foam core and the elastomeric outerlayer, wherein in being engineered, the foam core provides structuraland compressive support necessary to maintain a shape of the contouredstiffened composite panel during automated composite fiber placement aswell as autoclave curing, and wherein being engineered further comprisesthe elastomeric outer layer having a substantially constant thickness;enclosing said base composite layer and said at least one stiffeningelement in a vacuum bag and curing the base composite layer and the atleast one stiffening element; and removing the one-piece resilientmandrel body from said stiffening element cavity of said at least onestiffening element following said curing.
 2. The method of claim 1wherein said tooling surface comprises a generally concave contour. 3.The method of claim 1 wherein the one-piece resilient mandrel bodycomprises a generally triangular cross-section.
 4. The method of claim 1wherein the one-piece resilient mandrel body comprises a generallytrapezoidal cross-section.
 5. The method of claim 1 wherein theelastomeric outer layer comprises rubber.
 6. The method of claim 1,wherein removing includes deforming the elastomeric mandrel outer layer.7. A method for fabricating a contoured stiffened composite panel for anaircraft structure, comprising: providing a tooling surface having agenerally concave contour; placing a base composite layer on saidtooling surface; placing at least one stiffening element having astiffening element cavity on said base composite layer, the stiffeningelement comprising at least a stiffening element base surface and astiffening element side surface; providing a composite mandrel, saidcomposite mandrel comprising a generally triangular or trapezoidalcross-section and including a resilient foam mandrel one-piece corecoextensive with an elastic rubber mandrel outer layer disposed outsidesaid resilient foam mandrel one-piece core, the composite mandrelcomprising a composite mandrel base surface and a composite mandrel sidesurface, wherein a cross-sectional area and type of foam used for theresilient foam mandrel one-piece core is engineered to impartcompression compliance under autoclave pressure to offset a combinedthermal expansion behavior of the foam core and the elastomeric outerlayer during curing, wherein in being engineered, the foam core providesstructural and compressive support necessary to maintain a shape of thecontoured stiffened composite panel during automated composite fiberplacement as well as autoclave curing, and wherein being engineeredfurther comprises forming the elastic rubber mandrel outer layer to havea substantially constant thickness; inserting said composite mandrel insaid stiffening element cavity of said at least one stiffening elementsuch that the composite mandrel base surface contacts the stiffeningelement base surface and the composite mandrel side surface contacts thestiffening element side surface; enclosing said base composite layer andsaid at least one stiffening element in a vacuum bag and curing, duringcuring the composite mandrel base surface maintaining contact with thestiffening element base surface and the composite mandrel side surfacemaintaining contact with the stiffening element side surface so as toprevent collapse of the stiffening element; and removing said compositemandrel, after curing and without further heating, from said stiffeningelement cavity of said at least one stiffening element, the removingincluding deforming the elastic rubber mandrel outer layer so as toreduce an effort of removal.
 8. The method of claim 7 wherein thecomposite mandrel comprises a generally triangular cross-section.
 9. Themethod of claim 7 wherein the composite mandrel comprises a generallytrapezoidal cross-section.
 10. A method for fabricating a compositepanel with a stringer, comprising: placing a base composite layer on atooling surface, the tooling surface having a generally concave contour;placing a stiffening element having a cavity on the base compositelayer, the stiffening element comprising at least a stiffening elementbase surface and a stiffening element side surface; inserting acomposite mandrel in the cavity of the stiffening element, the compositemandrel comprising a core comprising a foam and an outer layercomprising an elastomeric material, the composite mandrel comprising acomposite mandrel base surface and a composite mandrel side surface, thecomposite mandrel substantially filling the cavity of the stiffeningelement such that the composite mandrel base surface contacts thestiffening element base surface and the composite mandrel side surfacecontacts the stiffening element side surface, wherein a cross-sectionalarea and type of foam used for the core is engineered to impartcompression compliance under autoclave pressure to offset a combinedthermal expansion behavior of the core and the outer layer, wherein inbeing engineered, the core provides structural and compressive supportnecessary to maintain a shape of the stringer during automated compositefiber placement as well as autoclave curing, and wherein beingengineered further comprises forming the outer layer to have asubstantially constant thickness; enclosing the base composite layer andthe stiffening element in a vacuum bag; curing the base composite layerand the stiffening element so as to form the composite panel with thestringer during curing the composite mandrel substantially filling thecavity of the stiffening element such that the composite mandrel basesurface contacts the stiffening element base surface and the compositemandrel side surface contacts the stiffening element side surface; andremoving the unitary composite mandrel from the cavity of the stiffeningelement, the removing including deforming the elastomeric material ofthe outer layer so as to reduce an effort associated with the removing.11. The method of claim 10 wherein the outer layer of the compositemandrel comprises an elastic rubber mandrel outer layer.
 12. The methodof claim 10, wherein the composite mandrel comprises a generallytriangular cross-section.
 13. The method of claim 10, wherein thecomposite mandrel comprises a generally trapezoidal cross-section.